A gas turbine engine typically includes a turbomachinery core having a high pressure compressor, combustor, and high pressure turbine in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. The high pressure compressor includes annular arrays (“rows”) of stationary vanes that direct air entering the engine into downstream, rotating blades of the compressor. Collectively one row of compressor vanes and one row of compressor blades make up a “stage” of the compressor. Similarly, the high pressure turbine includes annular rows of stationary nozzle vanes that direct the gases exiting the combustor into downstream, rotating blades of the turbine. Collectively one row of nozzle vanes and one row of turbine blades make up a “stage” of the turbine. Typically, both the compressor and turbine include a plurality of successive stages. Gas turbine engines, particularly aircraft engines, require a high degree of periodic maintenance. For example, engine components may exhibit wear during operation and often require repairs to restore its original dimensions and geometry.
Thermal spray coating repairs are often performed on engine components for dimensional build-up. However, the resulting repair material has debited properties due to oxidation and/or porosity as part of thermal spray processes. In particularly damaged components, such as those formed from nickel-based superalloys, thermal spray coating repair processes have not been successful for relatively thick coatings. That is, current repair processes, such as thermal spray coating, have not been very successful in depositing layers beyond about 1.5 mm thickness with properties similar to original base material.
As such, an improved method for repairing a nickel-based superalloy component is needed, particularly when requiring a coating of greater than 1.5 mm.